Thermally efficient tooling for composite component manufacturing

ABSTRACT

A method and apparatus for manufacturing composite components. A tool is present for use in manufacturing composite components. The tool comprises an encapsulation layer having a shape, an insulation layer on the encapsulation layer, and an isolation layer on the insulation layer. The isolation layer has an outer surface capable of contacting a composite material laid up on the outer surface. The insulation layer is capable of insulating the encapsulation layer from heat applied to the composite material. The encapsulation layer is capable of maintaining a shape with the composite material laid up on the isolation layer during a curing process to form a composite component from the composite material.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of and claims the benefit of priorityto U.S. patent application Ser. No. 13/645,052, filed Oct. 4, 2012, nowU.S. Pat. No. 8,834,773, which is a divisional of U.S. patentapplication Ser. No. 12/022,263, filed Jan. 30, 2008, now U.S. Pat. No.8,337,192, the entire contents of which are incorporated herein byreference.

BACKGROUND INFORMATION

1. Field

The present disclosure relates generally to composite components and inparticular to a method and apparatus for manufacturing compositecomponents. Still more particularly the present disclosure relates to amethod, apparatus, for manufacturing a composite component using a tool.

2. Background

Aircraft are being designed and manufactured with greater and greaterpercentages of composite materials. Some aircraft may have more thanfifty percent of their primary structure made from composite materials.Composite materials are used in aircraft to decrease the weight of theaircraft. This decreased weight improves performance features, such aspayload capacities and fuel efficiencies. Further, composite materialsprovide longer service life for various components in an aircraft.

Composite materials are tough, light-weight materials, created bycombining two or more dissimilar components. For example, a compositemay include fibers and resins. The fibers and resins are combined andcured to form a composite material.

Further, by using composite materials, portions of an aircraft may becreated in larger pieces or sections. For example, a fuselage in anaircraft may be created in cylindrical sections that may be put togetherto form the fuselage of the aircraft. Other examples include, forexample, without limitation, wing sections joined to form a wing,stabilizer sections joined to form a stabilizer, a stiffener, a fairing,a control surface, a skin, a skin section, a door, a strut, and atubular structure.

Currently, many composites in a manufactured aircraft require anautoclave to cure the composite components. An autoclave is a heatsource that provides both heat and pressure. Composite resins typicallyneed an elevated temperature to achieve a chemical reaction that allowsthese resins to flow and cure. Pressure is typically applied toconsolidate the materials in the part during resin flow. The temperaturetypically used is usually greater than 150 degrees Fahrenheit (typicallyabout 350 degrees Fahrenheit) with the pressure greater than oneatmosphere.

Further, in manufacturing composite components, the materials typicallyare formed using a mold. These molds also are referred to as tools. Atool has sufficient rigidity to maintain the desired shape for thecomposite component when the composite materials are placed onto thetools. A tool may be metallic or non-metallic in composition to providerigidity for supporting the composite materials.

With large components, a large autoclave is needed to encompass thecomponent and the tool for processing. In some cases, these largecomponents may be, for example, twelve to twenty feet in diameter andthey weigh tons.

As a result, composite materials have been developed which requirereduced heat and pressure to cure. Further, the amount of heat needed tocure these types of composite materials are typically at temperaturesless than around 250 degrees Fahrenheit. Pressures are provided byvacuum and an oven and/or heated molds used as the heat source forchemical reactions.

SUMMARY

The different advantageous embodiments provide a method and apparatusfor manufacturing composite components. In one advantageous embodiment,a tool is present for use in manufacturing composite components. Thetool comprises an encapsulation layer having a shape, an insulationlayer on the encapsulation layer, and an isolation layer on theinsulation layer. The isolation layer has an outer surface capable ofcontacting a composite material laid up on the outer surface whilemaintaining the final part shape. The insulation layer is capable ofinsulating the isolation layer from taking away heat applied to thecomposite material. The encapsulation layer is capable of maintainingthe shape with the composite material laid up on the isolation layerduring a curing process to form a composite component from the compositematerial.

In another advantageous embodiment, an apparatus comprises anencapsulation layer having a shape and an insulation layer located overthe encapsulation layer, wherein the insulation layer is capable ofinsulating the encapsulation layer from heat applied to a compositematerial and wherein the encapsulation layer is capable of maintainingthe insulation layer shape with the composite material laid up over theinsulation layer during a curing process to form a composite component.

In still another advantageous embodiment, a method is present formanufacturing a composite component. A composite material is placed on atool comprising an encapsulation layer that maintains a shape for thecomposite component; an insulation layer on the encapsulation layer; andan isolation layer on the insulation layer, wherein the isolation layerhas an outer surface in contact with the composite material laid up onthe outer surface, wherein the insulation layer insulates theencapsulation layer from heat applied to the composite material, andwherein the encapsulation layer maintains the shape with the compositematerial laid up on the isolation layer during a curing process to formthe composite component from the composite material. The compositematerial on the outer surface is cured to form the composite component.

The features, functions, and advantages can be achieved independently invarious embodiments of the present disclosure or may be combined in yetother embodiments in which further details can be seen with reference tothe following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the advantageousembodiments are set forth in the appended claims. The advantageousembodiments, however, as well as a preferred mode of use, furtherobjectives and advantages thereof, will best be understood by referenceto the following detailed description of an advantageous embodiment ofthe present disclosure when read in conjunction with the accompanyingdrawings, wherein:

FIG. 1 is a diagram illustrating an aircraft manufacturing and servicemethod in which an advantageous embodiment may be implemented;

FIG. 2 is a diagram of an aircraft in accordance with an advantageousembodiment;

FIG. 3 is a diagram of a composite component manufacturing system inaccordance with an advantageous embodiment;

FIG. 4 is a diagram illustrating examples of tools in accordance with anadvantageous embodiment;

FIG. 5 is a more detailed diagram illustrating examples of tools inaccordance with an advantageous embodiment;

FIG. 6 is a diagram illustrating another example of a tool in accordancewith an advantageous embodiment;

FIG. 7 is a diagram illustrating a cross-section of a tool in accordancewith an advantageous embodiment;

FIG. 8 is a flowchart of a process for creating a thermally efficienttool in accordance with an advantageous embodiment;

FIG. 9 is a flowchart of a process for manufacturing a compositecomponent in accordance with an advantageous embodiment; and

FIG. 10 is a graph illustrating time needed to heat a compositecomponent in accordance with an advantageous embodiment.

DETAILED DESCRIPTION

Referring more particularly to the drawings, embodiments of thedisclosure may be described in the context of the aircraft manufacturingand service method 100 as shown in FIG. 1 and aircraft 200 as shown inFIG. 2. Turning first to FIG. 1, a diagram illustrating an aircraftmanufacturing and service method is depicted in accordance with anadvantageous embodiment. During pre-production, exemplary aircraftmanufacturing and service method 100 may include specification anddesign 102 of aircraft 200 in FIG. 2 and material procurement 104.During production, component and subassembly manufacturing 106 andsystem integration 108 of aircraft 200 in FIG. 2 takes place.Thereafter, aircraft 200 in FIG. 2 may go through certification anddelivery 110 in order to be placed in service 112. While in service by acustomer, aircraft 200 in FIG. 2 is scheduled for routine maintenanceand service 114, which may include modification, reconfiguration,refurbishment, and other maintenance or service.

Each of the processes of aircraft manufacturing and service method 100may be performed or carried out by a system integrator, a third party,and/or an operator. In these examples, the operator may be a customer.For the purposes of this description, a system integrator may include,without limitation, any number of aircraft manufacturers andmajor-system subcontractors; a third party may include, withoutlimitation, any number of venders, subcontractors, and suppliers; and anoperator may be an airline, leasing company, military entity, serviceorganization, and so on.

With reference now to FIG. 2, a diagram of an aircraft is depicted inwhich an advantageous embodiment may be implemented. In this example,aircraft 200 is produced by aircraft manufacturing and service method100 in FIG. 1 and may include airframe 202 with a plurality of systems204 and interior 206. Examples of systems 204 include one or more ofpropulsion system 208, electrical system 210, hydraulic system 212, andenvironmental system 214. Any number of other systems may be included.Although an aerospace example is shown, different advantageousembodiments may be applied to other industries, such as the automotiveindustry.

Apparatus and methods embodied herein may be employed during any one ormore of the stages of aircraft manufacturing and service method 100 inFIG. 1. For example, components or subassemblies produced in componentand subassembly manufacturing 106 in FIG. 1 may be fabricated ormanufactured in a manner similar to components or subassemblies producedwhile aircraft 200 is in service 112 in FIG. 1.

Also, one or more apparatus embodiments, method embodiments, or acombination thereof may be utilized during production stages, such ascomponent and subassembly manufacturing 106 and system integration 108in FIG. 1, for example, without limitation, by substantially expeditingthe assembly of or reducing the cost of aircraft 200. Similarly, one ormore of apparatus embodiments, method embodiments, or a combinationthereof may be utilized while aircraft 200 is in service 112 or duringmaintenance and service 114 in FIG. 1.

The different advantageous embodiments recognize that with largercomposite components, the larger tools on which these components sithave additional mass that absorbs heat when curing composite materials.With the large size of tools, a large thermal mass is typically present.This thermal mass of the tool may absorb the heat generated by the heatsource away from the composite material. With the tool acting as a heatsink, the composite material takes more time to reach the desiredtemperature.

As a result, the amount of time needed to reach the appropriatetemperature to cure a composite material increases. This increasedheating time is also referred to as thermal lag and increases as themass of the tool. With curing systems using lower temperatures appliedby ovens or localized heat, the amount of time needed to cure thesetypes of composite components is greater as compared to those cured inautoclaves.

The problem increases in these types of curing systems because with lessheat being applied, more time is needed to cure the composite materialswith the lower temperatures. This problem is present even with localizedheating systems applied to the composite side because the tool acts as aheat sink and pulls heat away from the composite material.

Therefore, the different advantageous embodiments recognize thatimprovements to the current system for manufacturing compositecomponents are needed to decrease the amount of time needed tomanufacture components.

The different advantageous embodiments provide a method and apparatusfor fabricating composite components. The different advantageousembodiments recognize that removing or minimizing the tool as a heatsink may reduce the time needed to heat a composite material to theneeded temperature to cure that material.

As a result, in the different advantageous embodiments, a tool may beused to fabricate composite components in which the tool includes anencapsulation layer, an insulation layer, and an isolation layer. Theencapsulation layer is capable of maintaining the isolation and/orinsulation layer shape while curing a load from a composite materiallaid up on the tool to yield a proper shaped composite component.

The insulation layer is formed on or over the encapsulation layer andthe isolation layer is formed on or over the insulation layer. Theisolation layer has an outer surface that is capable of contacting thecomposite material laid up on the outer surface. The insulation layer iscapable of insulating the isolation layer and the encapsulation layerfrom heat transfer when heat is applied to the composite material. Insome embodiments, other materials or layers may be present between theselayers. These other layers may have various functions such as, forexample, binding two layers to each other.

Thus, less heat is conducted away from the composite material to theisolation layer with the use of an insulation layer. The isolation layermay be formed from an insulator or a thermal conductor. When a thermalconductor is used, the isolation layer is configured such that heatapplied to the composite material and the tool is distributed by theisolation layer to other areas on which the composite material is laidup on the outer surface. In this manner, the isolation layer also mayprovide a function in which more uniformed heating occurs.

With reference now to FIG. 3, a diagram of a composite componentmanufacturing system is depicted in accordance with an advantageousembodiment. Composite component manufacturing system 300 includescomposite material layup unit 302, tool 304, and heat source 306.Composite component manufacturing system 300 may be used to manufacturevarious types of composite components. For example, composite componentsmay be manufactured for aircraft. Examples of composite components mayinclude, for example, sections of a fuselage, an airframe, skin panels,and other suitable components.

Composite material layup unit 302 may place or layup composite material308 onto tool 304. Composite material layup unit 302 may take variousforms. For example, composite material layup unit 302 may be hand layupor automated tape layup machine or system, such as, for example, a M.Torres layup, which is a tape lay up machine available from M. Torres.Another example is Access-Atlas which is a composite working machineavailable from Forest-Liné. Yet another example of a machine that may beused is an Ingersoll Automated Tape Lamination Machine (ATLM), which isavailable from Ingersoll Machine Tools, Inc. Another example is anAutomated Fiber Placement Machine (AFPM). Of course, any type of systemor machine that may layup composite materials, such as tape, fabric,and/or any other suitable material may be used to implement compositematerial layup unit 302.

In these examples, tool 304 has a configuration that reduces the thermallag that may be encountered during curing of composite material 308.Tool 304 may take various forms, for example, tool 304 may be an innermold line or an outer mold line. Tool 304 may have less thermal massthan other tools that may be used to layup composite materials. Withless thermal mass, tool 304 may heat up to the desired temperature morequickly than currently used tools. Further, tool 304 also may bestructured to function less like a heat sink and more like an insulatorwith respect to the portions of tool 304 that contact composite material308.

As depicted, tool 304 includes isolation layer 310, insulation layer312, and encapsulation layer 314. In these examples, isolation layer 310has outer surface 316, which is capable of contacting composite material308. In other words, composite material 308 is laid up on outer surface316 of isolation layer 310. Isolation layer 310 may be a low thermaldiffusivity material or a high thermal diffusivity material. In otherwords, isolation layer 310 may be formed from either a thermal insulatoror a thermal conductor, depending on the particular implementation.

Thermal diffusivity may be calculated as follows:A=k/ρc _(p)Where a=Thermal Diffusivity; k=Thermal Conductivity; ρ=Density; andc_(p)=Specific Heat Capacity. When selecting materials for use of athermal insulator, such as that for isolation layer 310 or insulationlayer 312, a thermal diffusivity range of around less than 2.5E-06m²/sec. In selecting a material for a thermal conductor, the materialmay be selected as having a thermal diffusivity greater than around2.5E-06m²/sec.

Insulation layer 312 is made of materials having low thermal conductionor diffusivity in these examples. This layer may be made from variousmaterials and have various shapes. For example, a honeycombed shape, orfoam may be used. Additional examples of materials that may be used arediscussed below.

Encapsulation layer 314 provides the rigidity for tool 304 and maintainsthe shape of tool 304 while curing composite material 308 laid up onouter surface 316. Encapsulation layer 314 may be made from either aninsulator, conductor, depending on the particular implementation.Insulation layer 312 may be used to prevent or reduce heat applied tothe composite material from conducting through tool 304 to encapsulationlayer 314. When heat is not conducted away from composite material 308,the time needed to heat composite material 308 to a temperature neededto cure composite material 308 may be reduced.

Heat source 306 may heat composite material 308 laid up on outer surface316 of tool 304. By curing composite material 308 on tool 304, compositematerial 308 maintains the shape of tool 304. Heat source 306 may takevarious forms. For example, heat source 306 may be generated by oven 318or autoclave 320. Oven 318 provides heat while autoclave 320 providesheat and pressure to composite material 308 and tool 304.

Of course, other types of heat sources may be used, depending on theparticular implementation. For example, in addition to thermal curing asprovided by oven 318 and autoclave 320, other types of curing processesmay be employed. As another example, an electron beam system may be usedto cure composite material 308 to form the composite component. Heatsource 306 may be implemented using any currently available heat source.For example, an anchor autoclave from Anchor Autoclave Systems may beused or a composite curing autoclave from Taricco Corporation may beused.

With the use of tool 304, the amount of time needed to bring compositematerial 308 to the desired temperature for curing may be reduced. Thisreduction of time may occur because of the reduced thermal mass of tool304. With the configuration of tool 304 in these examples, the amount ofheat conducted away from composite material 308 to other layers of tool304 is reduced. When isolation layer 310 is thermally conductive, theheat is redirected to other portions of composite material 308 ratherthan being conducted through other layers of tool 304.

Further, with less thermally conductive materials within tool 304 ascompared to currently used tools, the amount of heat drawn into orneeded to heat tool 304 to the temperature needed to allow the compositematerial to be cured may be reduced.

Additionally, the weight of tool 304 also may be reduced lesseningstructural concerns for floors on which composite componentmanufacturing system 300 may be located. As a result, faster cure timesin oven 318 and autoclave 320 may be achieved. These faster cure timesmay result in reduced manufacturing costs, reduced work in process, andincreased part quantities using tool 304 with the existing manufacturingsystems.

Composite component manufacturing system 300 is depicted as one mannerin which a manufacturing system may be implemented. Of course, othercomponents may be present in addition to or in place of the onesillustrated in FIG. 3. For example, automated handling systems,composite material sources, computers, or other control mechanisms alsomay be present within composite material manufacturing system 300. Asanother example, additional members of composite material layup unitsand heat sources may be present within composite component manufacturingsystem 300.

In some advantageous embodiments, tool 304 may include insulation layer322 located below encapsulation layer 314. Insulation layer 322 mayfurther decrease the time needed to cure composite material 308 byproviding further insulation to tool 304 when encapsulation layer 314 isa thermal conductor, such as metal.

Turning now to FIGS. 4 and 5, diagrams illustrating examples of toolsare depicted in accordance with an advantageous embodiment. Tool 304from FIG. 3 may take various forms. As can be seen in this example, wing400 is a composite wing laid up using various tools. For example, bondjig 402, filler blocks 404, 406, 408, 410, 412, 414, 416, 418, 420, 422,and 424 may be implemented using tools, such as tool 304 in FIG. 3.Additionally, filler blocks 426, 428, 430, 432, 436, 440, 446, 448, 450,452, 454, 456, 458, 460, 462, 464, 466, and 468 may be implemented usinga tool, such as tool 304 in FIG. 3.

FIG. 5 is a more detailed example of section 469 in FIG. 4. This sectionprovides a more detailed illustration of some of the tools in FIG. 4,such as bond jig 402, filler block 404, filler block 406, filler block428, and filler block 430.

These types of tools may be used to layup composite materials forvarious components, such as stiffeners and skin panels in wing 400.Although the particular types of forms for a tool are depicted in FIGS.4 and 5, the different advantageous embodiments may be applied to anytype of tool used to create composite components from compositematerials using processes such as, for example, cure processes, cocureprocesses, co-bond processes, and bond processes.

Turning now to FIG. 6, a diagram illustrating another example of a toolis depicted in accordance with an advantageous embodiment. In thisexample, tool 600 has composite material 602 laid up on tool 600 for acomposite part. Tool 600 is located on stand 604. Tool 600 is comprisedof materials having a cross-section or layer similar to tool 304 in FIG.3. As can be seen, tool 600 provides the shape for the component that isformed from composite material 602.

With reference now to FIG. 7, a diagram illustrating a cross-section ofa tool is depicted in accordance with an advantageous embodiment. Inthis example, cross-section 700 is an example of a cross-section fromtool 600 in FIG. 6 taken along lines A-A. Layer 702 is an encapsulationlayer, such as encapsulation layer 314 in FIG. 3. Layer 704 provides acover for layer 702, which is an example of insulation layer 312 in FIG.3. Layer 704 is formed between layer 702 and layer 706. Layer 706 is anexample of an isolation layer, such as isolation layer 310 in FIG. 3.

Layer 706 has outer surface 708 which is a surface on which a compositematerial, such as layer 710, may be laid up or placed to form acomposite component. Layer 710 is laid on outer surface 708. Layer 706may be formed on layer 704 and may be made from different materials,such as an insulator or a conductor. When in the form of an insulator,layer 706 may be made from low thermal conduction materials, such aspolymers, adhesives, films, elastomers, organic fibers, and inorganicfibers. When polymers are used, the polymers may be filled or unfilled.

Of course, any other suitable materials may be used for layer 706depending on the implementation. Examples of inorganic materialsinclude, for example, glass, Kevlar, and ceramic fibers with polymermatrix. Layer 706 may be made of a different material from layer 704and/or may have a different structure. When using a thermal insulator, alow thermal conduction material may be selected. In these examples, alow thermal conduction material may have a thermal diffusivity of lessthan around 2.5E-06m²/sec.

When layer 706 takes the form of a thermal conductor, layer 706 may bemade from various materials, such as, for example, metals, thermallyconductive polymers, thermally conductive elastomers, or thermallyconductive organic materials. Examples of metals that may be usedinclude, for example, steel, aluminum, copper, and silver and alloysthereof. When a thermal conductor is used, a material having a highthermal conduction may be selected. This type of material may be onethat has a thermal diffusivity that is greater than around2.5E-06m²/sec.

With metals, solid metals, screens, or felts made of metals may be used.Further, layer 706 may be made from a combination of the differentinsulative or conductive materials. In these examples, layer 706 may beless than around 0.250 inches thick. When a thermal conductor isemployed, the durability of layer 706 may be increased as compared tousing a thermal insulator for layer 706.

Layer 704 may be made from various materials that have thermalinsulation properties. Layer 704 may use a material having a thermaldiffusivity of less than around 2.5E-06m²/sec. Layer 704 may be selectedto provide thermal insulation for layer 702 and/or layer 706 whilehaving a low mass or density to reduce weight. Layer 704 may be solid,porous, and/or include cavities. For example, layer 704 may take theform of foams, honeycombs, or fibers.

Further, layer 704 may comprise of at least one of foam, honeycombs, asheet, and fibers in a polymer matrix. At least one of an item meansthat layer 704 may be made from foam, honeycombs, a sheet, foam and asheet, foam and honeycombs, fibers in a polymer matrix, or some othercombination of the items listed.

When foam is used, the foam may be made from organic or inorganicthermal insulators. The honeycombs may be made from materials such as,for example, metallic materials, non-metallic materials, thermalelastomer materials, and/or paper. The fibrous insulation materials maybe, for example, fiberglass. An example of an organic material is balsawood. In the depicted examples, layer 704 may have a thickness fromaround 0.250 inches to around 12 inches.

Layer 702 covers insulation layer 704 and provides a shape for the tooland an ability to control the shape of the composite part formed fromthe composite material in layer 710. In other words, layer 702 providesstructural integrity to the different layers for handling and stabilityduring the manufacturing of composite components to maintain the shapedesired for the composite component.

Layer 702 may be made from various materials, such as organic materialsor inorganic materials. The form of layer 702 may be a solid sheet ofmaterial, such as, for example, organic, inorganic, and/or metallicmaterials. As another example, metallic fibers within a polymer matrixmay be used. Of course, any suitable form of material may be used thatis capable of providing sufficient structural integrity to a tool foruse in forming composite components. As depicted, layer 702 may have athickness of around less than 0.250 inches.

In this manner, the configuration of layers 706, 704, and 702 provide astructure for creating a thermally efficient tool, such as tool 304 inFIG. 3. In these examples, a thermally efficient tool is a tool thatreduces the amount of heat that the tool absorbs in a manner thatreduces the amount of time needed to heat a composite component forcuring as compared to currently used tools.

Although FIG. 7 depicts particular thermal diffusivity ranges, thicknessfor materials, types of materials, and form of materials, these examplesare presented only for purposes of illustrating some of the advantageousembodiments. Other advantageous embodiments may use other thermaldiffusivity levels and other ranges or thicknesses, depending on theparticular implementation.

With reference now to FIG. 8, a flowchart of a process for creating athermally efficient tool is depicted in accordance with an advantageousembodiment. The process illustrated in FIG. 8 may be used to create atool, such as tool 304 in FIG. 3.

The process begins by selecting materials for the tool (operation 800).These materials may be a combination of insulators and thermallyconductive materials as described above. Thereafter, the process formsan encapsulation layer (operation 802). Next, an insulation layer isformed on the encapsulation layer (operation 804). The process thenforms an isolation layer on the insulation layer (operation 806), withthe process terminating thereafter.

Of course, the construction of the different layers for a tool may varydepending on the particular embodiment. In some cases, the insulationlayer may be formed first with the encapsulation layer and the isolationlayer being placed on either side of the insulation layer. In somecases, the isolation layer may be formed first with the insulation layerformed on the isolation layer and the encapsulation layer formed on theinsulation layer. In some cases, the isolation layer and the insulationlayer and the encapsulation layer may be independently formed and thenbrought together to form the tool.

Turning now to FIG. 9, a flowchart of a process for manufacturing acomposite component is depicted in accordance with an advantageousembodiment. The process illustrated in FIG. 9 may be implemented using acomposite component manufacturing system, such as composite componentmanufacturing system 300 in FIG. 3.

The process begins by identifying a desired composite component(operation 900). Thereafter, a tool is selected for the identifiedcomposite component (operation 902). The composite material is placed onthe isolation layer of the tool to form an uncured composite component(operation 904).

Thereafter, the uncured composite component laid up on the tool isheated to cure the composite material to form the composite component(operation 906), with the process terminating thereafter.

Turning now to FIG. 10, a graph illustrating time needed to heat acomposite component is depicted in accordance with an advantageousembodiment. Graph 1000 illustrates the time needed to heat tools toaround 150 Fahrenheit degrees in an oven is depicted in accordance withan advantageous embodiment.

Line 1002 identifies the temperature in an oven. This line indicates thetime needed for the oven to heat to various temperatures, such as 150degrees. Line 1004 illustrates time needed for a tool constructedaccording to the advantageous embodiments to heat to 150 degrees. Usinga tool in accordance with an advantageous embodiment, the tool reaches150 degrees at one hundred and thirty minutes, which is around fiftyfive minutes after the over reaches 150 degrees.

Line 1006 illustrates time needed for a convention tool to be heated to150 degrees. The aluminum tool represented by line 1006 reaches 150degrees after one hundred and ninety five minutes, which is one hundredand twenty five minutes after the oven reached 150 degrees. As can beseen, the time needed for the conventional tool to heat to theappropriate temperature takes a longer amount of time as compared to atool according to the advantageous embodiments. Line 1008 depicts thetime needed to heat another tool that is commonly used for manufacturingcomposite components. The tool in line 1008 is made out of Invar®, whichis generically referred to as 36FeNi. As can be seen, this tool does notreach 150 degrees within two hundred and ninety minutes.

Thus, the different advantageous embodiments provide a method andapparatus for creating composite components. In the differentadvantageous embodiments, a tool having an insulation layer and anencapsulation layer is used. Composite materials may be placed on theinsulation layer in which the insulation layer prevents or reduces heatfrom conducting from the composite material to the encapsulation layer.As a result, thermal drive may be reduced in this manner.

In other advantageous embodiments, an isolation layer may be used inwhich the isolation layer is formed on the insulation layer and contactsthe composite material. This isolation layer may be an insulator or athermal conductor, depending on the particular implementation. Whentaking the form of a thermal conductor, the isolation layer may spreadout the heat applied to the composite material. As a result, if moreheat is applied in one section of the composite component, othersections may heat more evenly through the use of the isolation layer ina thermally conductive form.

In this manner, the amount of time needed to heat a composite materialto the temperature needed to cure the material may be reduced by using atool in accordance with an advantageous embodiment. As a result, theamount of time needed to create or manufacture a composite componentalso may be reduced. Although the different advantageous embodimentshave been described with respect to manufacturing composite componentsfor an aircraft, the different advantageous embodiments may be appliedto manufacturing composite components for other types of apparatus. Forexample, the different advantageous embodiments may be used tomanufacture composite components for spacecraft, submarines, ships,cars, trucks, manufacturing facilities, office buildings, and otherstructures.

The description of the different advantageous embodiments has beenpresented for purposes of illustration and description, and is notintended to be exhaustive or limited to the embodiments in the formdisclosed. Many modifications and variations will be apparent to thoseof ordinary skill in the art. Further, different advantageousembodiments may provide different advantages as compared to otheradvantageous embodiments. The embodiment or embodiments selected arechosen and described in order to best explain the principles of theembodiments, the practical application, and to enable others of ordinaryskill in the art to understand the disclosure for various embodimentswith various modifications as are suited to the particular usecontemplated.

What is claimed is:
 1. A method for manufacturing a composite component,the method comprising: providing a tool having an encapsulation layersandwiched between two insulation layers, and an isolation layer;placing a composite material in contact with an outer surface of theisolation layer; and curing the composite material on the outer surfaceto form the composite component from the composite material, wherein theinsulation layer insulates the encapsulation layer from heat applied tothe composite material; and placing the tool and the composite materialin an autoclave, wherein the tool reaches substantially 150 degreesFahrenheit no more than approximately 55 minutes after the autoclavereaches 150 degrees Fahrenheit; wherein the encapsulation layer providesrigidity to the tool and maintains a shape of the composite componentduring curing.
 2. The method of claim 1, wherein the curing stepcomprises: curing the composite material on the outer surface in theautoclave; wherein the tool is placed entirely in the autoclave.
 3. Themethod of claim 1, wherein the curing step comprises: curing thecomposite material on the outer surface in an oven.
 4. The method ofclaim 1, wherein the composite component is an aircraft part.
 5. Themethod of claim 4, wherein the aircraft part is selected from one of acylindrical section of a fuselage, a wing panel, and a section of astabilizer.
 6. The method of claim 1, wherein the insulation layer has athermal diffusivity less than about 2.5E-06m²/sec.
 7. The method ofclaim 1, wherein the insulation layer comprises a thermal insulationlayer material selected from one of organic insulating material,inorganic insulating material, a polymer, an adhesive, an elastomer, andglass.
 8. The method of claim 1, wherein the insulation layer comprisesat least one of foam and fibers.
 9. The method of claim 1, wherein theisolation layer comprises a thermal conducting material selected from ametal, a thermally conductive elastomer, a thermally conductive polymer,a thermally conductive organic material, and a metallic fiber.
 10. Themethod of claim 4, wherein the aircraft part is a stiffener.
 11. Themethod of claim 4, wherein the aircraft part is a fairing.
 12. Themethod of claim 4, wherein the aircraft part is a control surface. 13.The method of claim 4, wherein the aircraft part is a skin.
 14. Themethod of claim 4, wherein the aircraft part is a skin section.
 15. Themethod of claim 4, wherein the aircraft part is a door.
 16. The methodof claim 4, wherein the aircraft part is a strut.
 17. The method ofclaim 4, wherein the aircraft part is a tubular structure.
 18. Themethod of claim 1, wherein the insulation layer comprises honeycombs.19. The method of claim 1, wherein the insulation layer comprisessheets.
 20. The method of claim 1, wherein the insulation layercomprises a ceramic.